This invention concerns fuel injector assemblies for gas turbine engines.
There is a continuing need, driven by environmental concerns and governmental regulations, for improving the efficiency of and decreasing the emissions from gas turbine engines of the type utilised to power jet aircraft, marine vessels or generate electricity. Particularly there is a continuing drive to reduce nitrous oxide (NOx) emissions.
Advanced gas turbine combustors must meet these requirements for lower NO emissions under conditions in which the control of NO generation is very challenging. For example, the goal for the Ultra Efficient Engine Technology (UEET) gas turbine combustor research being done by NASA is a 70 percent reduction in NO emissions and a 15 percent improvement in fuel efficiency compared to ICAO 1996 standards technology. Realisation of the fuel efficiency objectives will require an overall cycle pressure ratio as high as 60 to 1 and a peak cycle temperature of 1600° C. or greater. The severe combustor pressure and temperature conditions required for improved fuel efficiency make the NOx emissions goal much more difficult to achieve.
Conventional fuel injectors that seek to address this issue have concentrically arranged pilot and main injectors with the main injector surrounding the pilot injector. However, conventional injector arrangements have several operational disadvantages, including for example, flame stability and re-light characteristics, the potential for excessive combustor dynamics or pressure fluctuations caused by combustor instability. Combustion instability occurs when the heat release couples with combustor acoustics such that random pressure perturbations in the combustor are amplified into larger pressure oscillations. These large pressure oscillations, having amplitudes of about 1-5% of the combustor pressure, can have catastrophic consequences and thus must be reduced or eliminated.
The invention seeks to provide an improved injector that addresses these and other problems.